An axial gas turbine engine includes a compressor section, a combustion section, and a turbine section. Disposed within the turbine section are alternating rows of rotatable airfoil blades and static vanes. As hot combustion gases pass through the turbine section, the airfoil blades are rotatably driven, turning a shaft and thereby providing shaft work for driving the compressor section and other auxiliary systems. The higher the gas temperature, the more work that can be extracted in the turbine section. During operation, the airfoils are constantly in contact with the hot working gases causing thermal stresses in the airfoils which effect the structural integrity and fatigue life of the airfoil. In an effort to increase the turbine section operating temperature, nickel or cobalt base superalloy materials are used to produce the turbine airfoil blades and vanes. Such materials maintain mechanical strength at high temperatures. However, even using such materials, it is necessary that the airfoil blades and vanes be cooled to maintain the structural integrity and fatigue life of the airfoil.
Numerous attempts have been made to provide internal cooling in airfoil structures. For example, in U.S. Pat. No. 3,171,631, issued to Aspinwall, titled "Turbine Blade", cooling air is flowed to a cavity between the suction sidewall and the pressure sidewall of an airfoil and diverted to various locations in the cavity by the use of turning pedestals or vanes. Another example is found in U.S. Pat. No. 3,533,712, to Kercher, titled "Cooled Vane Structure or High Temperature Turbines", where the use of serpentine passages extending throughout the cavity in the blade provides cooling to different portions of the airfoil. In U.S. Pat. No. 4,073,599, issued to Allen et al., titled "Hollow Turbine Blade Tip Closure", intricate cooling passages are coupled with other techniques to cool the airfoil. For example, the leading edge region in Allen et al. is cooled by impingement cooling followed by the discharge of the cooling air through a spanwise extending passage in the leading edge region of the blade.
In particular, small radius, high rotor speed engines require turbine blades which have highly twisted airfoils with a large variation of leading edge angle. A highly twisted airfoil has a high ratio of tip radius to root radius which provides a large change in airflow turning angle (camber) from root to tip, particularly in the leading edge area. While such a highly twisted leading edge has aerodynamic advantages, such a structure imposes severe restrictions on the design of the internal cooling structures required to obtain optimum leading edge cooling. In order to optimally cool the leading edge of such a blade, impingement holes must be incorporated internally which follow relatively precisely the leading edge angle. Most attempts to incorporate such impingement holes have been unsuccessful due to the difficulty in forming core dies which can accurately and consistently produce cores having the proper twist. Consequently, a need has arisen to provide a cooled, highly twisted airfoil which includes a structure for optimally cooling the leading edge region of the airfoil while minimizing processing time and reducing costs.